The present invention relates generally to gas turbine engines, and, more specifically, to turbine nozzles therein.
In a gas turbine engine, air is pressurized in a compressor and mixed with fuel in a combustor for generating hot combustion gases. The combustion gases are discharged from the combustor through a first stage turbine nozzle that channels the combustion gases into a row of turbine rotor blades which extract energy therefrom for powering the compressor.
The high pressure turbine (HPT) may have one or more turbine stages and is typically followed by a multistage low pressure turbine (LPT) that extracts additional energy from the combustion gases for powering an upstream fan in the typical turbofan aircraft engine configuration.
Since the first stage turbine nozzle first receives the high temperature combustion gases from the combustor it is subject to an extremely hostile operating environment that affects the useful life thereof. The nozzle components are typically formed from superalloys having enhanced strength at the experienced elevated temperatures of operation for maximizing useful life.
The turbine nozzle is subject to various pressure and thermal loads during operation which also effect corresponding stresses in the various components which stresses also affect nozzle life.
Since the nozzle thermally expands as it is heated by the combustion gases, and correspondingly thermally contracts as its temperature is reduced during the various operating cycles of the engine, substantial thermal loads and stresses are created in the nozzle. The thermal stresses therefore cycle in magnitude with the periodic operating cycles of the engine and its nozzle.
Accordingly, the life of the turbine nozzle itself is measured in operating cycles and is dependent upon the specific design of the turbine nozzle.
For example, typical turbine nozzles in large turbofan engines are circumferentially segmented into one or more vane segments to interrupt the circumferential continuity of the annular outer and inner bands which integrally support the corresponding turbine nozzle vanes therebetween.
Fully annular or unsegmented nozzle bands have increased strength and rigidity but correspondingly restrain expansion and contraction of the rigid nozzle vanes extending radially therebetween. Accordingly, significant thermal stresses are generated at the radial ends of the vanes where they integrally join their corresponding outer and inner bands.
Thermal restraint as well as structural rigidity are correspondingly reduced by circumferentially segmenting the nozzle bands, which correspondingly increases the complexity of the design by requiring suitable spline seals between the segmented bands.
A nozzle having a row of vane singlets has maximum segmentation of the bands with a single vane being integrally mounted to correspondingly short outer and inner band segments.
A nozzle having vane doublets includes two vanes integrally mounted in common band segments with correspondingly fewer segments around the perimeter.
And nozzle triplets are also known in which three vanes are integrally grouped to corresponding band segments for further reducing the segmentation of the bands.
However, as the number of vanes in each band segment increases, the significant problem of thermal restraint of the individual vanes also increases, with an associated increase in thermal stress where the vanes meet the integral bands.
Adding to the complexity of the design of modern turbine nozzles, is their mounting configuration in the engine itself. The nozzle is a fully annular assembly of components and must be suitably supported in the engine at the outlet end of the annular combustor with minimal thermal restraint that would otherwise add to the loads and stresses experienced by the nozzle.
Accordingly, the nozzle includes various flanges integrally formed in the inner and outer bands thereof, which flanges are used for mounting and sealing the nozzle in the engine, but which flanges also increase the structural rigidity of the nozzle and the corresponding thermal restraint.
The prior art is therefore replete with various forms of turbine nozzles having correspondingly different designs for use in correspondingly different gas turbine engines ranging in size and power from small to large for different aircraft and industrial applications.
In one conventional design of a small aircraft engine, a fully annular or unitary turbine nozzle is used without any circumferential segmentation of its outer and inner bands for reducing the structural complexity thereof, but at the expense of nozzle life.
The inner band includes a middle mounting flange, with the outer band including two pairs of circumferentially continuous flanges defining forward and aft annular grooves. Expansion seals in the form of split piston rings are trapped in the grooves and extend radially outwardly in sealing abutment with corresponding annular seal lands.
In this way, the unitary turbine nozzle is fixedly mounted in the engine from its inner band, with the outer band being allowed to freely expand and contract radially while the ring seals seal the pressurized gases.
However, operating experience has shown that this type of turbine nozzle has a finite useful life substantially less than that typically found for segmented turbine nozzles. And, in a present development program, it is desired to substantially increase the useful life of this type of nozzle for reducing maintenance outages and operating costs.
Accordingly, it is desired to provide a unitary turbine nozzle having reduced thermal stress for increasing useful life.